Methods and systems for determining a synthesized engine parameter

ABSTRACT

The present disclosure provides methods and systems for determining a synthesized engine parameter of a gas turbine engine. An initial model parameter is obtained from an onboard model associated with the gas turbine engine. A correction factor for the onboard model is determined by modifying a difference between the onboard model and an aero-thermal model of the gas turbine engine using first and second engine parameters and first and second operating conditions, wherein the first and second engine parameters are independent from one another over an operating envelope of the gas turbine engine. The initial model parameter is scaled by applying the correction factor thereto to obtain a corrected model parameter. The corrected model parameter is output as the synthesized engine parameter.

TECHNICAL FIELD

The present disclosure relates to gas turbine engines.

BACKGROUND

In a gas turbine engine, continuous inlet air is compressed, mixed withfuel in an inflammable proportion, and exposed to an ignition source toignite the mixture which then continues to burn to produce combustionproducts. The combustion of the air-fuel mixture can be used to powervarious mechanical components, which in turn can be used to producethrust.

Monitoring of various parameters within the engine during operationthereof can be of interest in assisting a control system or an operatorresponsible for the engine. Certain parameters may be difficult tomeasure using sensors, and are instead synthesized using a model of theengine. Although existing approaches for synthesizing engine parametersare suitable for their purposes, improvements remain desirable.

As such, there is room for improvement.

SUMMARY

In accordance with a broad aspect, there is provided a method fordetermining a synthesized engine parameter of a gas turbine engine. Aninitial model parameter is obtained from an onboard model associatedwith the gas turbine engine. A correction factor for the onboard modelis determined by modifying a difference between the onboard model and anaero-thermal model of the gas turbine engine using first and secondengine parameters and first and second operating conditions, wherein thefirst and second engine parameters are independent from one another overan operating envelope of the gas turbine engine. The initial modelparameter is scaled by applying the correction factor thereto to obtaina corrected model parameter. The corrected model parameter is output asthe synthesized engine parameter.

In accordance with another broad aspect, there is provided a system fordetermining a synthesized engine parameter of a gas turbine engine. Thesystem comprises a processing unit, and a non-transitorycomputer-readable medium. The computer-readable medium hascomputer-readable instructions stored thereon which are executable bythe processing unit for: obtaining an initial model parameter from anonboard model associated with the gas turbine engine; determining acorrection factor for the onboard model by modifying a differencebetween the onboard model and an aero-thermal model of the gas turbineengine using first and second engine parameters and first and secondoperating conditions, wherein the first and second engine parameters areindependent from one another over an operating envelope of the gasturbine engine; scaling the initial model parameter by applying thecorrection factor thereto to obtain a corrected model parameter; andoutputting the corrected model parameter as the synthesized engineparameter.

Features of the systems, devices, and methods described herein may beused in various combinations, in accordance with the embodimentsdescribed herein. In particular, any of the above features may be usedalone, together in any suitable combination, and/or in a variety ofarrangements, as appropriate.

BRIEF DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic cross-sectional view of an example gas turbineengine;

FIG. 2 is a block diagram of example engine models for the gas turbineengine of FIG. 1 ;

FIG. 3 is a block diagram of an example system for determining asynthesized engine parameter for the gas turbine engine of FIG. 1 ;

FIG. 4 is a graph illustrating an example correction curve associatedwith the system of FIG. 3 ;

FIG. 5 is a block diagram of an example computing device forimplementing part or all of the system of FIG. 3 ; and

FIGS. 6A-B are flowcharts illustrating example methods for determining asynthesized engine parameter of the gas turbine engine of FIG. 1 .

It will be noted that throughout the appended drawings, like featuresare identified by like reference numerals.

DETAILED DESCRIPTION

FIG. 1 illustrates a gas turbine engine 100 of a type provided for usein subsonic flight, generally comprising in serial flow communication, afan 112 through which ambient air is propelled toward an inlet 132, acompressor section 114 for pressurizing the air, a combustor 116 inwhich the compressed air is mixed with fuel and ignited for generatingan annular stream of hot combustion gases, and a turbine section 118 forextracting energy from the combustion gases, which exit via an exhaust136. High-pressure rotor(s) of the turbine section 118 (referred to as“HP turbine rotor(s) 120”) are drivingly engaged to high-pressurerotor(s) of the compressor section 114 (referred to as “HP compressorrotor(s) 122”) through a high-pressure shaft 124. Low-pressure rotor(s)of the turbine section 118 (referred to as “LP turbine rotor(s) 126”)are drivingly engaged to the fan rotor 112 and to low-pressure rotor(s)of the compressor section 114 (referred to as “LP compressor rotor(s)130”) through a low-pressure shaft 128 extending within thehigh-pressure shaft 124 and rotating independently therefrom. The gasturbine engine 100 can be provided with a nacelle 140 which defines abypass duct 142. Additionally, the gas turbine engine 100 includes ableed port 150. The bleed port 150 serves to extract air from the gaspath of the engine 100, which can be used as part of operations of thebroader system of which the engine 100 is part, for instance anaircraft.

Although illustrated as a turbofan engine, the gas turbine engine 100may alternatively be another type of engine, for example a turboshaftengine, also generally comprising in serial flow communication acompressor section, a combustor, and a turbine section, and an outputshaft through which power is transferred. A turboprop engine may alsoapply. Moreover, although the gas turbine engine 100 is illustrated as adual-spool engine, it should be understood that the techniques describedherein may also apply to single-spool engines. In addition, although theengine 100 is described herein for flight applications, it should beunderstood that other uses, such as industrial or the like, may apply.

Control of the operation of the engine 100 can be effected by one ormore control systems, for example an engine controller 110, which iscommunicatively coupled to the engine 100. The engine controller 110 canmodulate a fuel flow provided to the engine 100, the position andorientation of variable geometry mechanisms within the engine 100, ableed level of the engine 100, and the like, based on predeterminedschedules or algorithms. In some embodiments, the engine controller 110includes one or more FADEC(s), electronic engine controller(s) (EEC(s)),or the like, that are programmed to control the operation of the engine100. The operation of the engine 100 can be controlled by way of one ormore actuators, mechanical linkages, hydraulic systems, and the like.The engine controller 110 can be coupled to the actuators, mechanicallinkages, hydraulic systems, and the like, in any suitable fashion foreffecting control of the engine 100.

With additional reference to FIG. 2 , the behaviour of parts or all ofthe engine 100 may be simulated using one or more models, which aretypically generated by a manufacturer or designer of the engine 100. Themodels may be provided with various engine parameters and/or operatingconditions obtained by sensors associated with the engine 100, andsimulate the response of the engine 100 to the engine parameters and/oroperating conditions. Depending on the purpose and use case, modelshaving different levels of complexity and accuracy can be designed.Models having higher complexity may require additional computingresources to be properly implemented, including additional storage spaceand processing power. Conversely, models having lower complexity mayrequire reduced computing resources, and can be implemented on simplercomputing systems. In the embodiment illustrated in FIG. 2 , two modelsare developed: an aero-thermal model 210, and an onboard model 220.

The aero-thermal model 210 of the engine 100 is devised based on theengine 100. The aero-thermal model 210 provides a highly-accuratesimulation environment for the engine 100. The aero-thermal model 210would typically be employed by designers or manufacturers of the engine100, or of systems which use the engine 100, to determine how the enginewill behave in various scenarios. For instance, the aero-thermal model210 of the engine 100 may be used to devise control systems for theengine 100, to design aircraft which will employ the engine 100, and thelike.

The onboard model 220 is a simpler model than the aero-thermal model210, and therefore provides less accurate results when employed. In someembodiments, the onboard model 220 is designed for use by the enginecontroller 110 to simulate the behaviour of the engine 100 duringoperation thereof. For example, the engine controller 110 can use theonboard model 220 to simulate certain engine parameters of the engine100 which are not measured by sensors, or which cannot be synthesized inother fashions. The onboard model 220 may be developed based on theaero-thermal model 210, for instance by simplifying one or more aspectsof the aero-thermal model 210. For instance, one or more linearizationoperations are applied to the aero-thermal model 210 to generate theonboard model 220, which may involve a state-perturbation processassociated with particular inputs. Alternatively, the onboard model 220may be developed separately based on the operation of the engine 100.

As the onboard model 220 is less accurate than the aero-thermal model210, the engine parameters simulated by the engine controller 110 usingthe onboard model 220 may differ from those which would be produced bythe aero-thermal model using similar inputs. It may nevertheless bedesirable to synthesize engine parameters having a higher degree ofaccuracy using the onboard model 220. To this end, the presentdisclosure provides methods and systems for determining a synthesizedengine parameter for the engine 100 which may, in certain non-limitingembodiments, provide increased accuracy and/or precision for thesynthesized engine parameter. The methods and systems provided hereinmay make use of model parameters provided by the onboard model 220 andperform one or more correction steps in order to improve the accuracy ofthe model parameters.

With reference to FIG. 3 , there is illustrated a system 300 fordetermining a synthesized engine parameter, illustrated here ascorrected model parameter 330. The system is composed of a correctionmodule 310 and a database 320 which is communicatively coupled to thecorrection module 310. It should be noted that the system 300 may beimplemented by the engine controller 110, which also operates theonboard model 220. The correction module 310 is configured for producingthe corrected model parameter 330 based on an output from the onboardmodel 220, referred to herein as an initial model parameter 305. Theinitial model parameter 305 provided by the onboard model 220 can be anysuitable type of engine parameter, such as an engine temperature, anengine pressure, or any other value synthesized by the onboard model220. In some embodiments, the onboard model 220 is developed as a statevariable model, and the initial model parameter 305 can be any valuesynthesized by the state variable model.

The correction module 310 is configured for determining a correctionfactor for the onboard model 220. The correction factor can bedetermined in a variety of fashions, as will be described hereinbelow.In some embodiments, the correction module determines the correctionfactor based on a difference between the onboard model 220 and theaero-thermal model 210. The difference between the onboard model 220 andthe aero-thermal model 210 can be modified based on one or more engineparameters 302 of the engine 100, based on one or more operatingconditions 304 of the engine 100 or of a system of which the engine 100is a part, for instance an aircraft, or any suitable combinationthereof. For example, the correction module 310 can be provided withvalues for known differences between the onboard model 220 and theaero-thermal model 210, each associated with predetermined engineparameters 302 and/or operating conditions 304. The known differencescan be stored, for instance, in the database 320. The correction module310 may select one or more known differences which are associated withvalues of the engine parameters 302 and/or operating conditions 304 nearthe actual values therefor, and modify the known difference(s) to alignwith the actual values.

The correction factor is used to correct the initial model parameter305, thereby producing the corrected model parameter 330. The correctionfactor can be applied to the initial model parameter 305 in variousways. For example, the correction factor can include a multiplicativefactor, which is used to scale the initial model parameter 305 to obtainthe corrected model parameter 330. Alternatively, or in addition, thecorrection factor can include an additive factor, which is combined withthe initial model parameter 305 to obtain the corrected model parameter330. Other approaches are also considered.

Once the corrected model parameter 330 is determined, the correctionmodule 310 outputs the corrected model parameter 330. The correctionmodule 310 may output the corrected model parameter 330 to anotherportion of the engine controller 110, or to an outside component. Forinstance, in some embodiments in which the engine 100 forms part of anaircraft, the correction module 310 outputs the corrected modelparameter 330 to an avionics system and/or aircraft controllerassociated with the aircraft. In this fashion, the corrected modelparameter 330, which may be of a higher level of accuracy and/orprecision than the initial model parameter 305, can be provided to othersystems and/or devices for use thereby. As a result, the system 300 canbe used to improve the accuracy and/or precision of the onboard model220.

With reference to FIG. 4 , the difference between the onboard model 220and the aero-thermal model 210, as stored in the database 320 and usedby the correction module 310, may be provided in a variety of forms.FIG. 4 illustrates two example differences: a correction curve 400, anda linearized correction function 410. The correction curve 400 is basedon a collection of raw correction points 405, and is provided for apredetermined altitude of operation of the engine 100, and apredetermined rotor speed. For instance, the correction curve 400 may bea best-fit curve for the raw correction points 405. The raw correctionpoints 405 can be developed by a manufacturer or designer of the engine100, and is based on the difference between the onboard model 220 andthe aero-thermal model 210.

The linearized correction function 410 may represent a simplification ofthe correction curve 400 via one or more linearization processes. Forinstance, one or more splines are established to approximate thecorrection curve 400. One or more best-fit techniques may be performedto optimize the splines, thereby producing the linearized correctionfunction 410. In some cases, the linearized correction function 410 maybe defined by a collection of correction points, indicated by crosses at415, such that interpolating lines between the correction points 415produces the linearized correction function 410. The correction curve400 and the linearized correction function 410 detail the correctionfactor to be applied to the initial model parameter 305 as a function ofa pressure ratio for the engine 100.

In some embodiments, the database 320 can be provided with thecorrection curve 400 for use by the correction module 310. One or moreadditional correction curves 400 associated with different altitudes ofoperation, different rotor speeds, and/or with other different engineparameters 302 and/or operating conditions 304 can also be stored in thedatabase 320. However, in some other embodiments, the database 320 maynot be configured for storing the correction curve 400, or thecorrection module 310 may not be provided with the processingcapabilities for using the correction curve 400.

In some such embodiments, the correction curve 400 is not provided aspart of the system 300, and instead the correction points 415 areprovided. The correction points 415 are used to produce the linearizedcorrection function 410, which represents a simplified version of thecorrection curve 400: the inclusion of discrete points, rather than acurve or function, may result in storage space savings. Additionally,the use of the linearized correction function 410 based on thecorrection points 415 may also simplify calculations performed by thecorrection module 310. The database 320 can be provided with multiplecollections of correction points 415, each associated with differentaltitudes of operation, different rotor speeds, and/or with otherdifferent engine parameters 302 and/or operating conditions 304. In someother embodiments, one or more linearized correction function 410 may bestored in the database 320 and provided for use by the correction module310.

In one example implementation, multiple raw correction factors areobtained, based on evaluated differences between the onboard model 220and the aero-thermal model 210, across an operating space of the engine100. For example, differences between the onboard model 220 and theaero-thermal model 210 are evaluated at various altitudes of operation,temperatures, pressures, rotor speeds, and the like. In the case wherethe engine 100 is part of an aircraft, the evaluation of the differencebetween the onboard model 220 and the aero-thermal model 210 can also beperformed across different airspeeds. In some embodiments, theevaluation of the difference between the onboard model 220 and theaero-thermal model 210 at a predetermined set of engine parameters 302and operating conditions 304, and the associated raw correction factor,are determined based on the equation set:

E = S − A $C = {1 - \frac{E}{S}}$where A is an estimated parameter based on the aero-thermal model 210, Sis the estimated parameter based on the onboard model 220, E is theerror between the estimated parameters of both models 210, 220, and C isthe raw correction factor.

The raw correction factors are then grouped based on certain commoncharacteristics. For example, all the raw correction factors at a commontemperature, or for a common set of temperatures, are grouped. By way ofanother example, all the raw correction factors at a common altitude, orfor a common set of altitudes, are grouped. Similar approaches areconsidered for rotor speeds, or other engine parameters and/or operatingconditions. Within each group of correction factors, a non-linearfunction is used to fit the raw correction factors, for instance using aleast-squares regression method, or any other suitable regressiontechnique. The raw correction factors may be associated with certainschedules for the engine 100. The non-linear functions are thendiscretized using piecewise linear functions to approximate thenon-linear functions. The piecewise linear functions are then definedusing discrete coordinates, allowing them to be reconstructed usinginterpolation when needed by the correction module 310.

In some embodiments of the aforementioned example implementation, thegrouping of the raw correction factors is performed based on threetemperature categories: a first category associated with a firsttemperature dictated by the International Standard Atmosphere (ISA) fora given set of atmospheric parameters, a second category associated witha predetermined “extreme hot” temperature above the first temperature,and a third category associated with a predetermined “extreme cold”temperature below the first temperature. It should be understood thatother example implementations using more, or fewer, temperaturecategories are also considered. The grouping of the raw correctionfactors into these three groups can be used to generate variouscorrection curves, point collections, tables, or the like, for variousother parameters, including an altitude of operation of the aircraft.

In some embodiments of the aforementioned example implementation, one ormore engine parameters 302 and/or operating conditions 304 are selectedas main interpolation parameters for the difference between the onboardmodel 220 and the aero-thermal model 210. For instance, a particularpressure ratio of an inlet total pressure of the engine 100 to an inlettotal pressure basepoint, referred to herein as the inlet pressure ratioP_(2RATIO), can be used as a main interpolation parameter. In the caseof the inlet pressure ratio P_(2RATIO), the inlet total pressurebasepoint is a function of the inlet pressure of the engine 100 and theambient pressure in the vicinity of the engine 100, such that

$P_{2{RATIO}} = \frac{P_{2}}{P_{2{BASE}}}$$P_{2{BASE}} \propto \frac{P_{2}}{P_{AMB}}$where P₂ is the inlet total pressure of the engine 100, P_(2BASE) is theinlet total pressure basepoint, and P_(AMB) is the ambient pressure inthe vicinity of the engine 100. It should be noted that the inlet totalpressure P₂ is a function of both the ambient pressure P_(AMB) and theairspeed of the aircraft of which the engine 100 is a part.Additionally, the inlet total pressure basepoint P_(2BASE) is the valueof the inlet total pressure P₂ interpolated between basepoints of theonboard model 220. The basepoints of the onboard model 220 are definedas points within the operating envelope of the engine 100 where outputsof the onboard model 220 correspond to those of the aero-thermal model210. For example, a one-dimensional lookup table can be used tointerpolate values for the inlet total pressure basepoint P_(2BASE)between values of the inlet total pressure P₂ and the basepoints for theonboard model 220.

In some cases, multiple main interpolation parameters can be usedconcurrently. For instance, the inlet pressure ratio P_(2RATIO) is afirst main interpolation parameter, and a second interpolation parameteris a rotor speed of the engine 100. The rotor speed may be the speed ofthe high-pressure shaft 124 (sometimes referred to as N₂), the speed ofthe low-pressure shaft 128 (sometimes referred to as N₁), or anotherrotor speed associated with the engine 100. When multiple maininterpolation parameters are used, the selection of which engineparameters 302 and/or operating conditions 304 to use may be performedon the basis of the independence of the selected parameters. Forinstance, the inlet pressure ratio P_(2RATIO) and the low-pressure shaftrotor speed N₁ are independent of one another over a operating envelopeof the engine 100. That is to say, changes in the inlet pressure ratioP_(2RATIO) do not result in changes in the low-pressure shaft rotorspeed N₁, and vice-versa. In one example implementation, the inletpressure ratio P_(2RATIO) is used to schedule the correction factors fordifferent altitude groups across the operating envelope of the engine100, for average values of the low-pressure shaft rotor speed N₁, andfor the aforementioned temperature categories.

For example, the raw correction factors are grouped based on altitude ofoperation, low-pressure shaft rotor speed N₁, and temperature category.Best-fit non-linear curves as a function of the inlet pressure ratioP_(2RATIO) are generated based on the raw correction factors, which maycorrespond to the correction curve 400 of FIG. 4 . As discussedhereinabove, the curves can be discretized to piecewise linearapproximations. In some cases, the piecewise linear approximations arerepresented in a table format as coordinates which, when interpolated,reproduce the piecewise linear approximations. The different piecewiselinear approximations are each represented using a common number ofcoordinates in order to simplify their application.

With reference to FIG. 5 , there is illustrated an embodiment of acomputing device 510 for implementing part or all of the system 300described above. The computing device 510 can be used to perform part orall of the functions of the engine controller 110 of the engine 100. Insome embodiments, the engine controller 110 is composed only of thecomputing device 510. In some embodiments, the computing device 510 iswithin the engine controller 110 and cooperates with other hardwareand/or software components within the engine controller 110. In both ofthese cases, the engine controller 110 implements the functionality ofthe system 300. In some other embodiments, the computing device 510 isexternal to the engine controller 110 and interacts with the enginecontroller 110. In some other embodiments, some hardware and/or softwarecomponents are shared between the engine controller 110 and thecomputing device 510, without the computing device 510 being integral tothe engine controller 110. In both of these cases, the functionality ofthe system 300 can be implemented by the engine controller 110, by thecomputing device 510, or by a combination of both.

The computing device 510 comprises a processing unit 512 and a memory514 which has stored therein computer-executable instructions 516. Theprocessing unit 512 may comprise any suitable devices configured tocause a series of steps to be performed such that instructions 516, whenexecuted by the computing device 510 or other programmable apparatus,may cause the functions/acts/steps specified in the method 300 describedherein to be executed. The processing unit 512 may comprise, forexample, any type of general-purpose microprocessor or microcontroller,a digital signal processing (DSP) processor, a CPU, an integratedcircuit, a field programmable gate array (FPGA), a reconfigurableprocessor, other suitably programmed or programmable logic circuits, orany combination thereof.

The memory 514 may comprise any suitable known or other machine-readablestorage medium. The memory 514 may comprise non-transitory computerreadable storage medium, for example, but not limited to, an electronic,magnetic, optical, electromagnetic, infrared, or semiconductor system,apparatus, or device, or any suitable combination of the foregoing. Thememory 514 may include a suitable combination of any type of computermemory that is located either internally or externally to device, forexample random-access memory (RAM), read-only memory (ROM),electro-optical memory, magneto-optical memory, erasable programmableread-only memory (EPROM), and electrically-erasable programmableread-only memory (EEPROM), Ferroelectric RAM (FRAM) or the like. Memory514 may comprise any storage means (e.g., devices) suitable forretrievably storing machine-readable instructions 516 executable byprocessing unit 512.

It should be noted that the computing device 510 may be implemented aspart of a FADEC or other similar device, including an electronic enginecontrol (EEC), engine control unit (EUC), engine electronic controlsystem (EECS), an Aircraft Avionics System, and the like. In addition,it should be noted that the techniques described herein can be performedby a computing device 510 substantially in real-time.

With reference to FIG. 6A, the computing device 510 and/or the enginecontroller 110 may be configured for performing a method 600 fordetermining a synthesized engine parameter of a gas turbine engine, forexample the corrected model parameter 330 and the engine 100. The stepsdescribed in conjunction with the method 600 may also be implemented bythe system 300.

At step 610, the method 600 comprises obtaining an initial modelparameter from an onboard model associated with a gas turbine engine.For example, the initial model parameter 305 is obtained from theonboard model 220 of the engine 100. As noted hereinabove, the onboardmodel 220 may be implemented by the engine controller 110, which isassociated with the engine 100.

At step 620, the method 600 comprises determining a correction factorfor the onboard model 220 by modifying a difference between the onboardmodel 220 and an aero-thermal model of the engine 100, for instance theaero-thermal model 210. The difference may be modified using first andsecond engine parameters, which are independent from one another over anoperating envelope of the engine 100, and using first and secondoperating conditions. In embodiments in which the engine 100 is part ofan aircraft, the operating conditions include aircraft operatingconditions. For example, the engine parameters include the inletpressure ratio P_(2RATIO) and the low-pressure shaft rotor speed N₁. Inaddition, the operating conditions include the ambient temperature inthe vicinity of the engine 100, and the altitude of operation of theengine 100.

At step 630, the method 600 comprises scaling the initial modelparameter 305 by applying the correction factor, determined as part ofstep 620, to the initial model parameter 305 to obtain the correctedmodel parameter 330.

At step 640, the method 600 comprises outputting the corrected modelparameter 330 as the synthesized engine parameter. The corrected modelparameter 330 may be output by the correction module 310 of FIG. 3 tothe engine controller 110, or to one or more systems outside the enginecontroller 110. For instance, in embodiments in which the engine 100forms part of an aircraft, the corrected model parameter 330 may beoutput to an avionics system associated with the aircraft, or to anyother aircraft-level system, as appropriate.

It should be noted that the method 600 may be performed repeatedly formultiple different initial model parameters 305, and/or for multiplevalues of a common initial model parameter 305. For example, the initialmodel parameter 305 may be a pressure at a particular point within theengine 100, and the engine controller 110 may request that a correctedmodel parameter 330 for the pressure be produced at a predeterminedinterval, or at particular times, for instance in response to a request.In another example, an avionics system for an aircraft associated withthe engine 100 may request corrected model parameters for an enginetemperature and an engine pressure every few seconds. Other use casesare also considered.

With reference to FIG. 6B, an example approach for determining thecorrection factor for the onboard model 220, performed as part of step620 of the method 600, is illustrated. It should be noted that otherapproaches for determining the correction factor, as well as variationsto the approach described hereinbelow, are also considered.

At step 652, the step 620 comprises selecting three pairs of altitudecorrection tables based on an altitude of operation of the engine 100.The selection of the altitude correction tables in each pair isperformed by selecting a first altitude correction table which is abovethe altitude of operation of the engine 100, and by selecting a secondaltitude correction table which is below the altitude of operation. Thespacing between the first and second altitude correction tables may varybased on the altitude of operation, and the number of possible altitudecorrection tables available to the correction module 310. The altitudecorrection tables may be developed by a manufacturer or designer of theengine 100, and store coordinates much like the correction points 415described in conjunction with FIG. 4 . The altitude correction tablesare each associated with a respective temperature category: a firstaltitude correction table is associated with the ISA temperature, asecond is associated with a predetermined “extreme hot” temperatureabove the ISA temperature, and a third is associated with apredetermined “extreme cold” temperature below the ISA temperature.

At step 654, the step 620 comprises determining an inlet pressure ratiobased on an inlet total pressure and an inlet total pressure basepoint,for instance the inlet pressure ratio P_(2RATIO) described hereinabove.The inlet pressure ratio P_(2RATIO) may be determined based on measuredvalues for the inlet total pressure P₂, the ambient pressure P_(AMB),and the airspeed of the aircraft of which the engine 100 is a part.Other approaches are also considered.

At step 656, the step 620 comprises obtaining three pairs of preliminarycorrection factors by interpolating the three pairs of altitudecorrection tables with the inlet pressure ratio P_(2RATIO) and a rotorspeed of the engine 100, for instance the low-pressure shaft rotor speedN₁. The interpolation of the three pairs of altitude correction tablesconsists in mapping the inlet pressure ratio P_(2RATIO) and thelow-pressure shaft rotor speed N₁ to the three pairs of altitudecorrection tables to determine the three pairs of preliminary correctionfactors.

At step 658, the step 620 comprises obtaining an expected ambienttemperature by interpolating an altitude-ambient temperature correctiontable with the altitude of operation of the engine 100. Thealtitude-ambient temperature correction table can be developed by amanufacturer or designer of the engine 100, and is used to map an actualaltitude of operation of the engine 100 to an expected ambienttemperature.

At step 660, the step 620 comprises comparing an ambient temperature ina vicinity of the engine 100 to the expected ambient temperature. Atstep 662, the step 620 comprises selecting two of the three pairs ofpreliminary correction factors based on whether the ambient temperatureis above or below the expected ambient temperature. For example, whenthe ambient temperature is above the expected ambient temperature, thepair of preliminary correction factors associated with the ISAtemperature and the pair of preliminary correction factors associatedwith the predetermined “extreme hot” temperature above the ISAtemperature are selected. In another example, when the ambienttemperature is below the expected ambient temperature, the pair ofpreliminary correction factors associated with the ISA temperature andthe pair of preliminary correction factors associated with thepredetermined “extreme cold” temperature below the ISA temperature areselected.

At step 664, the step 620 comprises obtaining the correction factor byinterpolating the two selected pairs of preliminary correction factorswith the altitude of operation and the ambient temperature. Theinterpolation of the two selected pairs of preliminary correctionfactors can be performed similarly to the interpolation of the threepairs of altitude correction tables, discussed hereinabove in relationto step 656, or in any other suitable fashion. For example, thealtitude-based interpolation can be performed based on an altitude ratiobased on two altitude gaps. The first altitude gap is set as the valuebetween the altitude associated with the lower altitude correction tableand the altitude of operation of the engine, and the second altitude gapis set as the value between the lower and higher altitude correctiontables. Similarly, the temperature-based interpolation can be performedbased on a temperature ratio based on two temperature gaps. The firsttemperature gap is set as the value between the lower referencetemperature value (either the predetermined “extreme cold” temperatureor the ISA temperature) to the ambient temperature, and the secondtemperature gap is set as the value between the two selected referencetemperatures: either the predetermined “extreme cold” temperature to theISA temperature, or the ISA temperature to the predetermined “extremehot” temperature.

With additional reference to FIGS. 1 and 3 , in one example application,the method 600 of FIG. 6A, and in some cases the example implementationof the step 620 illustrated in FIG. 6B, are used to determine a bleedport pressure at the bleed port 150, referred to as P_(2.8). The bleedport pressure P_(2.8) may be of interest to an avionics system oranother aircraft-level system when the engine 100 is part of anaircraft.

In the example application for the bleed port pressure P_(2.8), aninitial bleed port pressure P_(2.8INIT) is obtained from the onboardmodel 220, which embodies the initial model parameter 305. Thecorrection module 310 uses the differences stored in the database 320 todetermine a correction factor to be applied to the initial bleed portpressure P_(2.8INIT) by modifying the difference using the engineparameters 302 and the operating conditions 304. For instance, thecorrection module 310 obtains the inlet pressure ratio P_(2RATIO), thelow-pressure shaft rotor speed N₁, the ambient temperature in thevicinity of the engine 100, and the altitude of operation of the engine100, which are used to modify the differences obtained from the database320.

Once the correction factor is determined by the correction module 310,the correction module scales the initial bleed port pressure P_(2.8INIT)using the correction factor. The correction module 310 thereby producesa corrected bleed port pressure, which can be output by the correctionmodule 310 as the actual bleed port pressure P_(2.8), which embodies thecorrected model parameter 330.

The methods and systems described herein may be implemented in a highlevel procedural or object oriented programming or scripting language,or a combination thereof, to communicate with or assist in the operationof a computer system, for example the computing device 510.Alternatively, the methods and systems described herein may beimplemented in assembly or machine language. The language may be acompiled or interpreted language. Program code for implementing themethods and systems described herein may be stored on a storage media ora device, for example a ROM, a magnetic disk, an optical disc, a flashdrive, or any other suitable storage media or device. The program codemay be readable by a general or special-purpose programmable computerfor configuring and operating the computer when the storage media ordevice is read by the computer to perform the procedures describedherein. Embodiments of the methods and systems described herein may alsobe considered to be implemented by way of a non-transitorycomputer-readable storage medium having a computer program storedthereon. The computer program may comprise computer-readableinstructions which cause a computer, or more specifically the processingunit 512 of the computing device 510, to operate in a specific andpredefined manner to perform the functions described herein.

Computer-executable instructions may be in many forms, including programmodules, executed by one or more computers or other devices. Generally,program modules include routines, programs, objects, components, datastructures, etc., that perform particular tasks or implement particularabstract data types. Typically the functionality of the program modulesmay be combined or distributed as desired in various embodiments.

The embodiments described in this document provide non-limiting examplesof possible implementations of the present technology. Upon review ofthe present disclosure, a person of ordinary skill in the art willrecognize that changes may be made to the embodiments described hereinwithout departing from the scope of the present technology. For example,additional numbers of altitude correction tables may be selected foradditional temperature categories. Yet further modifications could beimplemented by a person of ordinary skill in the art in view of thepresent disclosure, which modifications would be within the scope of thepresent technology.

The invention claimed is:
 1. A method for determining a synthesizedengine parameter of a gas turbine engine of an aircraft, comprising:generating an initial model parameter by a computing device running anonboard model; providing the initial model parameter to a correctionmodel implemented on the computing device; determining by the computingdevice, a correction factor for the onboard model by modifying adifference between the onboard model and an aero-thermal model of thegas turbine engine using first and second engine parameters and firstand second operating conditions, wherein the first and second engineparameters are independent from one another over an operating envelopeof the gas turbine engine, the determining of the correction factor forthe onboard model including selecting three pairs of altitude correctiontables based on an altitude of operation of the gas turbine engine, eachpair of altitude correction tables among the three pairs of altitudecorrection tables associated with a respective temperature category, andobtaining three pairs of preliminary correction factors by interpolatingthe three pairs of altitude correction tables with an inlet pressureratio and a rotor speed of the gas turbine engine; scaling the initialmodel parameter by the computing device, by applying the correctionfactor thereto to obtain a corrected model parameter; and outputting thecorrected model parameter from the computing device as the synthesizedengine parameter to another system comprising an avionics system and/oraircraft controller associated with the aircraft.
 2. The method of claim1, wherein determining the correction factor for the onboard modelcomprises: comparing an ambient temperature in a vicinity of the gasturbine engine to an expected ambient temperature based on the altitudeof operation; selecting two of the three pairs of preliminary correctionfactors based on whether the ambient temperature is above or below theexpected ambient temperature; obtaining the correction factor byinterpolating the two selected pairs of the three pairs of preliminarycompensation factors with the altitude of operation and the ambienttemperature.
 3. The method of claim 2, further comprising obtaining theexpected ambient temperature by interpolating an altitude-ambienttemperature correction table with the altitude of operation.
 4. Themethod of claim 1, wherein selecting the three pairs of altitudecorrection tables based on the altitude of operation comprises, for eachpair of altitude correction tables of the three pairs of altitudecorrection tables, identifying a first altitude correction tableassociated with an altitude value above the altitude of operation andidentifying a second altitude correction table associated with analtitude value below the altitude of operation.
 5. The method of claim1, further comprising determining the inlet pressure ratio based on aninlet total pressure and an inlet total pressure basepoint, wherein theinlet total pressure basepoint is based on a measured inlet pressure anda measured ambient pressure.
 6. The method of claim 1, furthercomprising generating the altitude correction tables based on theaero-thermal model of the gas turbine engine.
 7. The method of claim 1,wherein obtaining the correction factor comprises: determining a firstaltitude gap between an altitude of operation of the gas turbine engineand a first altitude associated with a first temperature category of theonboard model; determining a second altitude gap between the firstaltitude and a second altitude associated with a second temperaturecategory of the onboard model; and determining an altitude gap ratiobetween the first altitude gap and the second altitude gap.
 8. Themethod of claim 1, wherein obtaining the correction factor comprises:determining a first temperature gap between an ambient temperature in avicinity of the gas turbine engine and a first temperature associatedwith a first temperature category of the onboard model; determining asecond temperature gap between the first temperature and a secondtemperature associated with a second temperature category of the onboardmodel; and determining a temperature gap ratio between the firsttemperature gap and the second temperature gap.
 9. The method of claim1, further comprising obtaining an initial modeling signal as an outputof the onboard model, wherein the initial modeling signal is a bleedport pressure.
 10. A system for determining a synthesized engineparameter of a gas turbine engine, comprising: a processing unit; and anon-transitory computer-readable medium having stored thereoncomputer-readable instructions executable by the processing unit for:obtaining an initial model parameter from an onboard model associatedwith the gas turbine engine; determining a correction factor for theonboard model by modifying a difference between the onboard model and anaero-thermal model of the gas turbine engine using first and secondengine parameters and first and second operating conditions, wherein thefirst and second engine parameters are independent from one another overan operating envelope of the gas turbine engine, the determining of thecorrection factor for the onboard model including selecting three pairsof altitude correction tables based on an altitude of operation of thegas turbine engine, each pair of altitude correction tables among thethree pairs of altitude correction tables associated with a respectivetemperature category, and obtaining three pairs of preliminarycorrection factors by interpolating the three pairs of altitudecorrection tables with an inlet pressure ratio and a rotor speed of thegas turbine engine; scaling the initial model parameter by applying thecorrection factor thereto to obtain a corrected model parameter; andoutputting the corrected model parameter as the synthesized engineparameter.
 11. The system of claim 10, wherein determining thecorrection factor for the onboard model comprises: comparing an ambienttemperature in a vicinity of the gas turbine engine to an expectedambient temperature based on the altitude of operation; selecting two ofthe three pairs of preliminary correction factors based on whether theambient temperature is above or below the expected ambient temperature;obtaining the correction factor by interpolating the two selected pairsof the three pairs of preliminary compensation factors with the altitudeof operation and the ambient temperature.
 12. The system of claim 11,wherein the computer-readable instructions are further executable forobtaining the expected ambient temperature by interpolating analtitude-ambient temperature correction table with the altitude ofoperation.
 13. The system of claim 10, wherein selecting the three pairsof altitude correction tables based on the altitude of operationcomprises, for each pair of altitude correction tables of the threepairs of altitude correction tables, identifying a first altitudecorrection table associated with an altitude value above the altitude ofoperation and identifying a second altitude correction table associatedwith an altitude value below the altitude of operation.
 14. The systemof claim 10, wherein the computer-readable instructions are furtherexecutable for determining the inlet pressure ratio based on an inlettotal pressure and an inlet total pressure basepoint, wherein the inlettotal pressure basepoint is based on a measured inlet pressure and ameasured ambient pressure.
 15. The system of claim 10, wherein thecomputer-readable instructions are further executable for generating thealtitude correction tables based on the aero-thermal model of the gasturbine engine.
 16. The system of claim 10, wherein obtaining thecorrection factor comprises: determining a first altitude gap between analtitude of operation of the gas turbine engine and a first altitudeassociated with a first temperature category of the onboard model;determining a second altitude gap between the first altitude and asecond altitude associated with a second temperature category of theonboard model; and determining an altitude gap ratio between the firstaltitude gap and the second altitude gap.
 17. The system of claim 10,wherein obtaining the correction factor comprises: determining a firsttemperature gap between an ambient temperature in a vicinity of the gasturbine engine and a first temperature associated with a firsttemperature category of the onboard model; determining a secondtemperature gap between the first temperature and a second temperatureassociated with a second temperature category of the onboard model; anddetermining a temperature gap ratio between the first temperature gapand the second temperature gap.
 18. The system of claim 10, wherein thecomputer-readable instructions are further executable for obtaining aninitial modeling signal as an output of the onboard model, wherein theinitial modeling signal is a bleed port pressure.